Turbofan case and method of making

ABSTRACT

A casing for a gas turbine includes a construction providing improved structural efficiency. Improved load paths and means for transmitting loads in the engine case are disclosed.

CROSS-REFERENCE

This application is a division of application Ser. No. 10/883,987, filedJul. 6, 2004, now U.S. Pat. No. 7,266,941, which is a continuation inpart of application Ser. No. 10/628,556, filed Jul. 29, 2003, now U.S.Pat. No. 7,370,467.

FIELD OF THE INVENTION

This invention relates to gas turbine engines, and more particularly toa case for a turbofan engine.

BACKGROUND OF THE INVENTION

The general aviation market has recently been introduced to the “verysmall” turbofan engine (i.e. 2000 pounds thrust and less). Simplyscaling down larger conventional turbofan engines, however, presentsdifficulties due mainly to the disproportionate scaling of certainfactors, such as strength-to-weight and tolerances.

The engine case, such as that depicted in FIGS. 1 and 2, is subjected toasymmetric loading relative to the engine mounts, caused by loadsexerted through the bearings, such as engine thrust, foreign objectimpacts and blade-off events, and caused by inertia loads caused by theengine weight which of course must be supported. These asymmetric loadsresult in bending moments and shears which must be transmitted throughthe engine case to the engine mounts. The prior art generally relies onthick walled structures, such as cast engine case components (such as202 204, 206, 208, 210, 211 in FIG. 1), to react these bending momentsin plate bending. Plate bending, however, requires thicker-walledcasings to resist and carry bending forces without failure. In verysmall engines, however, thick casing become a significant component inoverall engine weight.

An alternate approach is shown in U.S. Pat. No. 4,132,069, whichprovides an integrally-webbed structure for transferring loads, and inparticular bending, through an engine and nacelle structure, so thatnacelle loads can be passed to the engine. The scheme, however, addscomponents to the engine, which reduces reliability, and increasesweight and cost. Improvement in engine case technology is thereforedesired.

SUMMARY OF THE INVENTION

It is therefore one object of the present invention to provideimprovements gas turbine engine case technology.

In accordance with one aspect of the present invention, there isprovided a casing for an aircraft turbofan bypass engine, the casingcomprising a case adapted to encircle the engine and having a pluralityof engine mounts thereon adapted to mount the engine to an aircraft; aninner hub adapted to support at least one bearing supporting a mainshaft of the engine, the inner hub supported inside the case by aplurality of struts extending between the inner hub to the case, thestruts defining a primary load path from the inner hub to the case; anda splitter supported intermediate the inner hub and case by the struts,the struts further defining a primary splitter load path from thesplitter to the case, the splitter adapted to divide an engine ingestedairflow between a core airflow passage and a bypass airflow passage ofthe engine, wherein the case has a semi-monocoque configurationincluding a plurality of ribs and a plurality of thin-walled shearpanels therebetween, the case thereby being adapted to balance externalloading applied to the casing by compressive and tensile forces in theribs to react balanced shear in the panels.

In accordance with another aspect of the present invention, there isprovided an aircraft bypass turbofan engine comprising an engine coreand a casing surrounding at least a portion of the engine core, thecasing including a plurality of hollow struts and a plurality ofadjoining members, the struts extending in a circumferential arraybetween an inner hub and the casing, each of the struts adjoined to atleast two circumferentially adjacent struts by at least one of themembers, the members each having two end portions each mounted to astrut side, each member comprising a hollow closed section, the closedsection at least partially closed by the strut sides and at least oneelement extending between adjacent struts, the element adapted by reasonof its alignment relative to the member and adjacent struts to transmita shear force into the struts when a torque is applied to the member.

In accordance with a further aspect of the present invention, there isprovided an aircraft bypass turbofan engine casing comprising a outerring portion, an inner hub portion and a plurality of hollow struts anda plurality of hollow torque box members, the outer ring portion havingat least one engine mount thereon for engine-supporting connection to anaircraft, the struts arranged in a circumferential array and extendingfrom the inner hub portion to the outer ring portion to mount the innerhub portion to the outer ring portion, the plurality of torque boxmembers arranged such that at least one extends between adjacent strutsin the array to thereby connect each strut to immediately adjacentstruts, the torque box members adapted to convert a torque appliedthereto into a shear force and transmit said shear force into thestruts.

In accordance with a further aspect of the present invention, there isprovided a load carrying apparatus for a aircraft bypass turbofanengine, the apparatus comprising an inner hub supporting at least onemain shaft bearing; an outer casing having at least one engine mount;and a hollow strut assembly including a plurality of struts extending inan circumferential array, the plurality of struts each extending from afirst end connected to the inner hub to a second end connected to theouter casing, the struts having sides facing immediately adjacent strutsin the array, the strut assembly including means for load sharingbetween adjacent struts, said means extending between adjacent strutsand connecting to an intermediate portion of each strut side.

In accordance with a further aspect of the present invention, there isprovided a load carrying apparatus for a aircraft bypass turbofanengine, the apparatus comprising an outer ring and an inner ringtogether defining at least one air flow passage therebetween, aplurality of hollow struts extending radially between the outer andinner rings across the passage in a circumferential array, and aplurality of hollow torque boxes, each torque box bonded withshear-transferring joints to an intermediate portion of adjacent struts,the outer ring having a plurality of engine mounts for mounting theengine to an aircraft, the torque boxes including a web member adaptedto transfer a torque applied to the torque box by an engine core mountedthereto into the strut as shear for engine core load transfer to theengine mounts.

In accordance with a further aspect of the present invention, there isprovided a casing for an aircraft bypass turbofan engine, the casingcomprising a case adapted to encircle the engine and having a pluralityof engine mounts thereon adapted to mount the engine to an aircraft; aninner hub adapted to support at least one bearing supporting a mainshaft of the engine, the inner hub supported relative to the case by aplurality of struts extending between the inner hub to the case, theinner hub having a semi-monocoque configuration including a plurality ofstiffeners and a plurality of thin-walled shear panels therebetween, theinner hub thereby being adapted to resolve external bending forcesapplied to the inner hub substantially as compressive and tensile forcesin the stiffeners and shear in the panels.

In accordance with a further aspect of the present invention, there isprovided a casing for an aircraft bypass turbofan engine, the casingcomprising a case adapted to encircle the engine and having a pluralityof engine mounts thereon adapted to mount the engine to an aircraft; aninner hub adapted to support at least one bearing supporting a mainshaft of the engine, the inner hub supported relative to the case by aplurality of struts extending between the inner hub to the case, theinner hub having a semi-monocoque configuration including a plurality ofstiffeners and a plurality of thin-walled shear panels therebetween, thestiffeners and panels configured to react external bending momentsapplied to the inner hub as compressive and tensile forces in thestiffeners and shear in the panels.

In accordance with a further aspect of the present invention, there isprovided a casing for an aircraft turbofan bypass engine, the casingcomprising a case adapted to encircle the engine and having a pluralityof engine mounts thereon adapted to mount the engine to an aircraft; aninner hub adapted to support at least one bearing supporting a mainshaft of the engine, the inner hub supported inside the case by aplurality of struts extending between the inner hub to the case, thestruts defining a primary load path from the inner hub to the case; anda splitter supported intermediate the inner hub and case by the struts,the struts further defining a primary splitter load path from thesplitter to the case, the splitter adapted to divide an engine ingestedairflow between a core airflow passage and a bypass airflow passage ofthe engine, wherein the struts include means in a trailing edge portionthereof for interrupting a load path between the splitter and inner hubto thereby inhibit the transfer of splitter loads to the inner hub.

In accordance with a further aspect of the present invention, there isprovided a shaft bearing support apparatus for a gas turbine engine, theapparatus comprising a bearing support member, a stop apparatus and astop surface, wherein the stop apparatus and stop surface are subject torelative deflection therebetween when a shaft supported by a bearingmounted to the bearing support member deflects in use, and wherein aclearance is provided between the stop apparatus and stop surface equalto a maximum desired magnitude of said relative deflection such thatcontact between the stop apparatus and the stop surface occurs when saidmaximum desired relative deflection occurs, the stop apparatus and thestop surface thereby being adapted to arrest deflection beyond saidmaximum desired relative deflection by reason of said contact.

In accordance with a further aspect of the present invention, there isprovided a casing for an aircraft turbofan bypass engine, the casingcomprising a case adapted to encircle the engine and having a pluralityof engine mounts thereon adapted to mount the engine to an aircraft; aplurality of struts extending between an engine structure and the case,the struts defining a primary load path from the engine structure to thecase for transfer of loading from engine structure to the engine mountson the case, the struts each having a centroidal axis defined along alocus of centroid positions for a plurality of strut sections along alength of the strut, wherein the engine mounts are positioned on thecase to substantially correspond with at one of said strut centroidalaxes to thereby minimize bending loads in the case as a result of loadstransferred by the struts to the engine mounts.

In accordance with a further aspect of the present invention, there isprovided a casing for an aircraft turbofan bypass engine, the casingcomprising a case adapted to encircle the engine and having a pluralityof engine mounts thereon adapted to mount the engine to an aircraft; aplurality of struts extending between an engine structure and the case,the struts defining at least one load path from the engine structure tothe case for transfer of loading from engine structure to the enginemounts on the case, wherein at least some struts are adapted toplastically deform in response to the application of a pre-selected loadalong said load path thereto thereby limiting load transfer from thestruts to the engine mounts by said struts to an amount below saidpre-selected load. Also disclosed is a method of providing such acasing.

Still other features and advantages of the present invention will bebetter understood with reference to the preferred embodiments describedhereinafter.

It should be noted that the terms of “integral”, “integrating” and“integrated” used throughout the text of this application and appendedclaims, are intended to mean items which are integrally joined such thatdisassembly (in a typical non-destructive sense) is not possible.

BRIEF DESCRIPTION OF THE DRAWINGS

Having thus generally described the nature of the present invention,reference will now be made to the accompanying drawings, showing by wayof illustration the preferred embodiments thereof, in which:

FIG. 1 is a simplified exploded perspective view of a conventional caseassembly of a turbofan engine;

FIG. 2 is a schematic cross-sectional view of a similar conventionalcase assembly;

FIG. 3 is a schematic cross-sectional view of a turbofan case accordingto the present invention;

FIG. 4 is a schematic partial cross-sectional view of the embodiment ofFIG. 3;

FIG. 5 is a exploded isometric view, with a portion cut away, of anintermediate portion of the assembly of FIG. 4;

FIG. 6 is an exploded isometric view of the assembly of FIG. 4,illustrating the assembly sequence of the intercase portion of FIG. 5;

FIG. 7 is an isometric front view of the intercase portion shown inFIGS. 5 and 6;

FIG. 8 is an isometric rear view of the intercase portion shown in FIGS.5-7;

FIG. 9 is an exploded and enlarged isometric front view of a portion ofan alternate embodiment of the intercase portion of the presentinvention;

FIG. 10 is an enlarged isometric front view of a cross-section of theassembled case of the present invention;

Amended FIG. 11 is an enlarged cross-sectional view of a portion of thepresent invention showing the fan exit vane installation;

Amended FIG. 12 is a somewhat schematic cross-sectional view showingassembly steps according to the present invention;

FIG. 13 is an enlarged view of a portion of FIG. 12;

FIG. 14 is a partial top view of the case of FIG. 13;

FIG. 15 is a rear view of the case of FIG. 12;

FIG. 16A is a schematic representation of the force transfer in thesplitter and strut of the case of FIG. 12 from a perspective similar tothat of FIG. 15;

FIG. 16B is a schematic representation of the force transfer in thesplitter and strut of the case of FIG. 12 from a perspective similar tothat of FIG. 15;

FIG. 16C is a schematic representation similar to that of FIG. 16A,showing an alternate configuration for the splitter;

FIG. 17 is a somewhat schematic top plan view of the inner hub of thecase of FIG. 12;

FIG. 18A is a cross-sectional view through the strut of FIG. 12;

New FIG. 18B shows a view of a prior art strut from a perspectivesimilar to that of FIG. 18A; and

New FIG. 19 is a somewhat schematic view of an alternate configurationfor the strut of FIG. 12.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

Referring to the drawings, beginning with FIG. 3, an exemplary turbofangas turbine engine 10 according to the present invention includes inserial flow communication about a longitudinal central axis 12, a fanassembly 13 having a plurality of circumferentially spaced fan blades14, a compressor section 16 having a plurality of circumferentiallyspaced low pressure compressor (LPC) blades 50 and high pressurecompressor (HPC) blades 51, a diffuser 18, a combustor 20, a highpressure turbine (HPT) 22, and a low pressure turbine (LPT) 24. LPT 24is connected to the fan assembly 13 by a first or low pressure (LP)shaft 26, and HPT 22 is connected to compressor assembly 16 by a secondor high pressure (HP) shaft 28. Fuel injecting means 30 are provided forinjecting fuel into the combustor 20

A generally tubular casing assembly 32 having a envelops the engine 10and thereby defines a main flow path 36 through the core of engine 10,extending from an inlet 34 to an exhaust outlet (not shown), and aby-pass flow path 37.

Referring to FIGS. 3, 4 and 6, the casing assembly 32 according to oneembodiment of the present invention includes a generally tubular fanportion or “case” 44, which houses the fan rotor assembly 13, agenerally tubular intercase or intermediate portion or “case” 46downstream of fan case 44 and a gas generator portion or “case” 52downstream of intermediate portion 46. The intermediate portion 46includes a compressor shroud 48 which encircles the blade tips of thecompressor assembly 16, and a bearing seat 58 for mounting the HP shaftbearing 59 thereto, as will be described further below.

With reference to FIGS. 5 and 6, gas generator portion 52, which is alsogenerally tubular in shape, is for housing the combustor 20 and perhapsHPT 22 or a section thereof. A generally tubular case turbine andexhaust case 54 is preferably modularly provided and mounted to (i.e.not integrated with) the aft end 107 of gas generator case 52 forhousing the LPT 24, and supporting an exhaust mixer assembly (notshown).

The engine 10 further includes a tubular bypass duct case 56, preferablymodularly provided and mounted to (i.e. not integrated with) theintermediate portion 46 of casing assembly 32. The tubular bypass ductcase 56 generally surrounds the gas generator portion 52 and is radiallyspaced apart therefrom, thereby defining a downstream section of thebypass 44 therebetween.

Rather than providing a prior art segmented case, in which the casecomponents are removably mounted to one another, the present inventionprovides a single-piece casing assembly 32 in which all casingcomponents are integrally attached to one another. Referring again toFIG. 3, fan case portion 44, intermediate case portion 46, compressorshroud portion 48, bearing mount 58 and gas generator portion 52 ofcasing assembly 32 are all integrally joined to one another, such as bywelding, or by other process such as integral fabrication, brazing orother methods of joining and bonding the components into one piece.Preferably, the bypass duct case 56 is not integrated with casing 32, inorder to provide convenience in assembly and maintenance of the engineassembly 10, and so rather is connected by bolting together matingflanges 60 and 62 which extend radially from the respective intermediateportion 46 and the bypass duct case 56. The turbine and exhaust case 54,as mentioned, is also preferably mounted to the aft end of the casing 32by, for example, bolting together mated flanges 64 and 66. The bypassduct 56 and the case 54 are shown by broken lines in FIG. 4 todistinguish them from other cases which are most preferably integratedto form the integral case of the present invention. Casing assembly 32can also integrally include the bypass and exhaust ducts, if desired.

The individual components of casing 32 are preferably made from onematerial, for example steel, although a combination of materials may beused (e.g. steel and Inconel, etc.) as long as the desired integralbonding technique (e.g. welding) permits such materials to be reliablybonded together. The individual portions of the casing are preferablymade separately, as will be described further below, which would permit,for example, a variety of processes and materials to be used.Optionally, the casing 32 may be formed integrally substantially in asingle operation, such as metal injection moulding.

Surprisingly, although the entire casing 32 of the present invention maybe made from a relatively heavy material such as steel, in very smallturbofan engines (i.e. preferably 2000 pounds thrust and less, morepreferably 1500 pounds thrust and less, and most preferably about 1000pounds thrust or less) the present invention provides unexpected andsignificant benefits which directly impact on engine SFC, as will now bedescribed.

Firstly, even though a heavier material is used throughout (e.g. steelversus, say, magnesium), the weight savings from reduced flange count issurprisingly significant. Even scaled-down flanges represent asignificant weight relative to the very small turbofan engine, and thusit has been found that their removal results in a disproportionateweight savings despite the addition of weight elsewhere in the casing,contrary to the teachings of the prior art. Therefore, contrary to theteachings of the prior art, it has been found that a segmented casepermitting the use of lighter materials is actually heavier in the verysmall turbofan range. A beneficial redistribution of weight is thereforeprovided by the present invention.

Secondly, the reduction of flange connections also beneficially reducestolerance stack-up by reducing the number of toleranced parts andconnections. Accordingly, for example by integrating the compressorbearing mount and compressor shroud into a single part, a significantlysmaller compressor blade tip clearance may be provided.

Thirdly, the reduction of thermally mismatched parts also permits asignificant simplification to the very small turbofan engine. In a firstaspect, the reduction of thermal mismatch improves the tolerances whichmust be left in connections. In a second aspect, by improving thermalmismatch within the casing 32, the interface with other systems, such asthe accessory gearbox (AGB) is greatly simplified.

In a second aspect of the present invention, a configuration for casing32 is disclosed which provides further benefits to the very smallturbofan. Referring to FIGS. 4 and 5, the structure of the intermediateportion 46 of casing 32 will now be described in more detail. Theintermediate portion 46 includes an outer ring 68 having a forward end70 and a rearward end 71 integrated with the radially outwardlyextending bypass duct flange 60. On the external surface of the outerring 68 are provided stiffening ribs 72, which reinforce the rigidity ofthe outer ring 68, and engine mounts 74 which also assist in thisregard. As can be seen in FIGS. 5 and 6, ribs 72 are arranged in agrid-like manner relative to one another and thereby divide outer ring68 into a plurality of panels 68B. A mounting support 82 on the outerring 68 is provided for operatively supporting the AGB tower shaft (notshown), and to provide further stiffness to ring 68. Also provided onthe outer ring 68 are attachment brackets 84 for attaching the AGB.Other services, such as oil tube inlet 83 and N1 probe boss 85, are alsoprovided.

The intermediate portion 46 of casing 32 also includes an inner hub 76which has a forward end 78 and a rearward end 80. The inner hub 76 ispositioned coaxially with the outer ring 68 and is supported within theouter ring 68 by a plurality of casing struts 40 which arecircumferentially spaced apart and extend radially outwardly andgenerally rearwardly from the inner hub 76 to the outer ring 68, as willbe described further below. The annular bearing seat 58 which receivesand supports preferably the HPC bearing 59 (see FIG. 3) is integrallyattached (for example, by welding, as described below) to the rearwardend 80 of the inner hub 76. A mounting flange 77 is also provided on theforward end 78 of the inner hub 76 (see FIGS. 4 and 5) for attaching aforward bearing housing (not shown) for the LP shaft bearings.

The intermediate portion 46 of casing 32 also includes a splitter 42,which includes an annular inner wall 85 and an annular outer wall 86extending axially and downstream relative to the air flow through engine10, divergent from an annular leading edge tip 88. A section of theannular bypass path 37 is thereby defined between the outer ring 68 andthe annular outer wall 86 of the splitter 42, while core flow path 36 isdefined between the annular inner wall 85 of the splitter 42 and theinner hub 76. An internal web 94 is provided within splitter 42, betweenthe inner and outer walls 85, 86, and affixed thereto, and preferablyalso affixed to struts 40, as will be described further below. Asdescribed previously, the compressor shroud 48, which is preferablythicker than the inner wall 85 of the splitter 42 to withstand thedemands of the compressed air flow, is integrated (for example bywelding, as described further below) to the inner wall 85.

A plurality of circumferentially spaced apart slots 90 extend generallyfrom near the annular tip 88 axially into the splitter 42, for receivingthe respective casing struts 40. A plurality of corresponding bosses 91and 93 are respectively provided in the inner hub 76 and the outer ring68 for attaching the casing struts 40.

A bleed valve housing 92 (see FIGS. 4 and 6) is preferably attached bywelding, to the annular outer wall 86 of the splitter 42 at its rearwardend, for securing bleed valve(s) (not shown) thereto. The intermediateportion 46 also bleed holes 96 defined in the outer wall 86 of thesplitter 42, for co-operation with an air bleed system (not shown).Bleed holes 96 are preferably made when fabricating the splitter 42.

Though when assembled it has the appearance of a prior art intercase,which is most typically cast, the present invention advantageouslypermits the individual components of intermediate portion 46 may be madein accordance with a variety of manufacturing processes. The preferredprocesses will now be described. Outer ring 68 and inner hub 76 aremachined from solid. Outer ring 68 is generally quite thin (i.e.sheet-metal-like) and, in conjunction with stiffener ribs 72, provideintercase portion 46 with a semi-monocoque construction which islightweight yet strong. Service attachments, such oil tube inlet 83 andN1 probe boss 85, are cast (or metal injection moulded, forged,machined, etc., as desired) and welded or brazed to outer ring, whileother “attachments” such as tower shaft support 82 are integrallymachined with the ring. Struts 40 are formed preferably in sheet metalhalves (though processes such as metal injection moulding, hydroforming,flow forming, casting, etc. may be used) and then integrally joined bywelding to provide a hollow configuration. One strut preferably receivesan AGB tower shaft (not shown), another the oil tube and N1 probe (notshown), and so on. The struts 40 are preferably welded to bosses 91 and93 and within slots 90, to thereby assemble outer ring 68, splitter 42and inner hub 70 to provide intercase portion 46 of casing 32.

Referring to FIG. 9, in an alternate embodiment, intercase portion 46may have struts 40 which have a configuration which provides a modifiedjoint with splitter 42 and outer ring 68, through the inclusion offlanged components 40A and 68A which may be welded to struts 40 andouter ring 68 respectively. Such flanged components may be provided tofacilitate stronger connection welds, etc. and thus this embodimentfurther illustrates the flexibility the present invention gives thedesigner.

The individual components are integrated together preferably by welding(or other integral joining technique of the general types alreadymentioned) to provide the integrated intermediate portion 46, and thisis preferably before integrating the intermediate portion 46 with theother portions of the casing 32 (i.e. fan portion 44, etc.). The detailsof the intermediate portion 46 may vary depending on various embodimentsused for various engine models.

Referring to FIGS. 4 and 6, the fan portion 44 includes an annularupstream section 98 encircling the fan blades 14 (see FIG. 3). Theupstream section 98 is preferably strong enough to ensure containment ofa blade-off incident, or incorporate an insert therefor (not shown). Thefan case 44 includes a downstream section 100 which extends from theupstream section 98 to a downstream edge 103. The downstream section 100incorporates slots 101 which locates and supports the outer end of fanexit vanes 38, as will be described below.

Referring to FIG. 10, the stator-less fan exit vanes 38 are slidinglyinserted preferably from outside the fan portion 44 and therefore slots101 are defined accordingly in the section 100 of the fan portion 44(see FIG. 6) and in the inner shroud 102. The fan exit vanes 38 arereleasably mounted between the section 100 of the fan portion 44 and theinner shroud 102 at the corresponding slots, and releasably retainedtherein by pliable compression-fit insert grommets 120 (see FIG. 11) andstraps 122.

Fan portion 44 may be flow-formed from one material, such as steel,nickel or inconel. Alternate fabrication or forming techniques may alsobe used, and one or more materials may be used.

The fan portion 44 is integrated into the intermediate portion 46 byintegrally joining, preferably by welding, the aft end 103 of fan caseportion 44 with the forward end 70 of the outer ring 68 of theintermediate portion 4 to thereby create an integral joint 130 (see FIG.4). The inner shroud 102 of the fan portion 44 is also attached to theinner hub 76 of the intermediate portion 46, preferably by welding at132. The inner shroud 102 and the fan exit vanes 38 are preferably notintegrated with the casing assembly 32, but rather are releasablymounted to the fan portion 44 as described above after the fan portion44 is integrated with the intermediate portion 46.

The gas generator case portion 52 of casing 32, includes an upstreamsection 104 and a substantially cylindrical downstream section 106 whichare integrated together, preferably by being fabricated in a singlemanufacturing process. An integral inner ring 108 is disposed within theupstream section 104 and is integrated, preferably by welding, with thegas generator case 52 at the forward end thereof. A mounting flange 110extends radially outwardly from the inner ring 108 at the inner edgethereof, for securing the diffuser 18 flange 110A and bleed valve 150thereto (see (FIGS. 3, 4 and 12). A number of openings 140 (see FIG. 6)are provided in the gas generator case 52 for receiving or mountingengine components of the gas generator portion, such as fuel injectingmeans 30, and so on, as will be understood by one skilled in the art.The downstream cylindrical section 106 has an aft end 107 which isintegrated with a radially outwardly extending mounting flange 112, forconnection with turbine and/or exhaust case 54. The gas generator case52 is integrated at the front end thereof with the aft end 89 of theannular outer wall 86 splitter 42 of the intermediate portion 46 at 134,also preferably by welding.

The fan portion 44, the intermediate portion 46 and the gas generatorportion 52 of casing 32 are thus fabricated separately, for example bymachining from solid, sheet metal fabrication, forging, casting,flow-forming, etc., depending on the design of each and the wishes ofthe designer. The separately fabricated cases are then integrallyattached preferably by welding. It is then preferable to finally machinethe interior portions of the integrated casing 32 prior to installationof rotor assemblies, in order to reduce any tolerance stack-up occurringduring casing 32 manufacture or assembly. This dramatically reduces thetolerance stack-up over prior art devices.

The way in which each portion is formed and the exact means by which theportions are attached are not critical to the invention, but rather maybe left to the designer's discretion. Therefore, the present inventionallows for flexibility in selection of manufacturing processes to meetthe designer's needs in providing an integrated case assembly for a verysmall turbofan engine. The present invention thereby permits a varietyof manufacturing techniques, notably among them fabrication techniquessuch as machining from solid, flow-forming and sheet metal construction,which are not available with prior art casing designs.

In yet another aspect of the present invention, the flexibility ofmanufacture permitted by the present invention permits the bearingmounts integrally provided in the case to be much simpler, in terms ofpart count, than prior art bearing mounts. Typical prior art gas turbineengines require complicated bearing mounts, including assemblies knownas “squirrel cages” to dampen vibrations caused by rotor imbalanceswhich inevitably result despite highly accurate machining processes. Inthe present invention however, bearing mounts such as bearing mount 58may be provided with an integrated flexibility, such that which is afunction of its material, configuration, stiffness, etc., such thatbearing mount 58 itself can be “tuned” during manufacture to therebyobviate the need for a squirrel cage. The bearing mount 58 is thusintegrally designed and provided to also perform a damping function toremove the need for separate squirrel cage assemblies. Since squirrelcages add weight, length and complexity to the engine, deleting thiscomponent is of course valuable and therefore yet another beneficialfeature of the present invention.

Referring now to FIGS. 5, 6 and 12, in a yet further aspect of thepresent invention, a method for assembling a turbofan engine will now bedescribed. Unlike the prior art, the present invention casing 32 ispreferably fully (or substantially) assembled before any rotating orother gas turbine components are assembled therein. Thus, the first stepis making and assembling the components of the casing assembly 32, asdescribed above. The next step, also described above, preferably is tomachine internal surfaces of the casing 32, such as surfaces relating tobearing mounts, compressor shrouds and similar surfaces, to remove anyaccumulated tolerance stack-up which would affect the efficientoperation of the engine. The next steps are to insert the fan rotorassembly 13 inside casing 32 (step not shown in the Figures), preferablythrough the inlet 34 of the casing assembly 32 and into the fan portion44, and to insert the bleed valve 150 and compressor assembly 16 intocasing 32, preferably through gas generator portion 52 (see FIG. 12).The diffuser 18, combustor 20, the turbine assemblies, and othercomponents are also inserted into casing 32, also preferably from theaft end of the gas generator portion 52. The assembly process of theengine 10 is then completed by further mounting the turbine and exhaustcase 54, the bypass duct 56, and other engine components in and to thecasing assembly 32. While the specific order of insertion and assemblyof these interior assemblies in casing may depend on preference or thedesign layout of engine 10, the present invention involves building thecore of engine 10 inside a completed or substantially completed casing32, thereby permitting an overall more efficient assembly technique forthe gas turbine engine.

The present method also advantageously provides a fast assembly of a gasturbine engine because no fixtures such as flange connections arerequired and therefore, less “final” assembly steps are required.

As mentioned, the present invention has particular application for usein so-called very small gas turbine engines, namely engines typically2000 pounds thrust and below for use in general aviation aircraftsometimes referred to as “personal” jet aircraft. This market representsa leading edge of gas turbine turbofan technology, wherein the limits ofscaling and cost-effective design and operation are challenged. Priorart small turbines, such as those used in missile engines are simplyunsuitable. Missile engines are invariably expensive to make and operate(owing to their military-heritage), and are designed for extremely shortoperational lives (a few hours) in which they are continuously operatedat full thrust. The very small turbofan as contemplated herein, however,must of course be operated intermittently at varying thrust levels (e.g.idle, taxi, take-off, climb, cruise, approach and landing) for thousandsof hours, not to mention be affordable and quiet to operate andenvironmentally friendly. Likewise, although microturbines are beginningto proliferate in the power generation field, this technology is alsolargely unsuitable since aircraft applications require extremelylightweight and reliable designs which are typically not found inindustrial microturbine designs. Accordingly, the present inventionrepresents an advance in the field of providing an affordable-to-operateturbofan to general aviation pilots.

The present invention permits a turbofan casing to be provided which, inthe very small turbofan size range, permits the overall weight of thecasing to be reduced over conventional larger designs. The weightreduction is due in part to the thin shell stiffened semi-monocoquedesign of the intermediate case section 46, which has anintegrally-stiffened thin shell construction which allows the designerto optimize the use of metal to thereby reduce weight. Referring againto FIGS. 5, 6 and 7, the thin “sheet” outer ring “panels” 68B arereinforced at specific locations by the ribs 72 and struts 40, and byengine mounts 74 and other similar features on the ring 68, to balanceexternal loading by compression and tension in the reinforcing membersreacting balanced shear in the “panels” 68B of the outer ring 68. Thisprovides a stable structure with a stiffness comparable to a caststructure more than 500% thicker. It is through this approach, combinedwith the simplicity of attachment, that the overall weight of the casingis significantly reduced.

Referring again to FIGS. 5 6 and 7, as described above, outer ring 68has a thin-walled semi-monocoque design includes a plurality of ribs 72extending axially and circumferentially about the outer ring 68 tothereby define a plurality of thin-shell panels 68B therebetween. Theaxial and circumferential arrangement of ribs 72 provides panels 68Bwith a generally rectangular shape and the ribs being more or lessparallel or perpendicular to one another. A partial top view of outerring 68 is shown in FIG. 14, showing ribs 72 and thin-shell panels 68B.

The splitter 42 separates core flow passage 36 from bypass flow passage37, and is supported by. Each strut 40 extends from a leading edge 40Ato trailing edge 40B, the trailing edge having a bent, kinked ordiscontinuous profile having an inner portion 40C and an outer portion40D joined by a bend or kink 40E. Each strut 40 extends from an innerend to an outer end (not indicated) to meet with and connect to bosses93 and 91, respectively, integrally provided on inner and outer rings.

Referring now to FIG. 13, the splitter 42 is joined to the strut 40 andincludes the internal web 94 (see also FIGS. 3-5) which co-operates withstruts 40 and splitter 42 to thereby define a plurality ofclosed-section hollow torque boxes 41 between adjacent struts 40 (seealso FIG. 15). In the example engine depicted in FIG. 15, therefore,since there are six struts there are six torque boxes 41 formedtherebetween. Struts 40, splitter 42 and web 94 are joined to oneanother by shear-transmitting joints (e.g. welded, brazed, or otherbonded joint, or have an integral construction and hence not be “joints”per se). The joints (indicated by 42A and 94A in FIG. 16 a) arepreferably strong enough provide the necessary shear connections toprevent deformation of the torque boxes under anticipated loadings, aswill be described below. These torque boxes provide the mechanism fortransferring the bending moments associated with the weight of theengine core transferred from the gas generator case to the splitter (seeFIGS. 3, 4, and 6, for example).

The splitter 42 preferably further includes a circumferential stiffeningring 43 slightly aft of torque box 41. Similarly, the inner hub 76preferably includes a pair of circumferential stiffening rings 76A, and76B, respectively, on an interior side thereof, and preferably axiallypositioned to correspond to the locations at which struts 40, boss 91meet inner hub 76. The Inner hub 76 supports the main low spool thrustbearings at bearings 57 and also includes a bearing attachment seat 58and a bearing bumper 58A, as will be described in more detail below.

Mounts 74 are preferably positioned relative to struts 40 such thatmounts 74 are substantially aligned with a centroidal axis “CA” (seeFIG. 13) of strut 40 to thereby significantly reduce any tendency forloads to cause strut bending relative to the mounts 74. The ‘centroidalaxis’ will be understood to mean a line passing through the centroids ofall axial sections of a strut 40 (i.e. will pass through the centroid ofany horizontal section of the strut 40, as viewed in FIG. 13).

As mentioned above, outer ring 68, which is a semi-monocoque structurecomposed of thin-shell shear panels 68B, and axial and circumferentialstiffeners 72, is thus analogous to conventional aircraft fuselageturned inside-out. The loads applied to the structure are reacted aseither tension or compression (depending on the direction of the sourceload) in the ribs 72, which are internally balanced by opposing shearsin the panels 68A. Stresses are thus shared amongst adjacent ribs 72,and bending forces are avoided by resolution to in-plane tensile andcompressive forces and shear. This manner of reacting loads in sheargives the intermediate case portion 46 a relatively high structuralefficiency and stiffness compared to a typical prior art cast enginecase. In the design described, engine mounts 74 and strut bosses 93 alsoact as tensile/compressive load bearing members communicating withadjacent shear panels. Loads thus enter the outer ring 68 via the struts40/bosses 93, and are passed through the semi-monocoque structure orribs and shear panels to the engine mounts 74, for ultimate transmissionto the aircraft. Since out-of-plane bending forces are resolved intoin-plane compressive/tensile loads, the think prior art case sectionsare not required as bending is no longer reacted merely by the casingsection in plate bending. The result is a casing which is significantlylighter than the prior art, particularly when high modulus materials areused, such as steel. Although the ribs & panel configuration shown inFIG. 14 is preferred, the grid need not be regular nor rectangular, butrather any effective configuration preferred by the designer may beused.

Similar to outer ring 68, inner hub 76 is also provided with asemi-monocoque structure, as follows. Stiffener rings 76A and 76B andstrut bosses 91 co-operate to divide the annular surface of hub 76 intoa plurality of thin-shell shear panels 76C which react to tensile orcompressive loads in rings 76A, 76B and strut bosses 91 as a shear inpanels 76C, as depicted in FIG. 17, to thereby balance the structure. Inthis manner, bending in the inner hub is minimized such that the panels76C may be substantially thinner than the prior art (e.g. the presentinvention may have panels of 0.050″ or less). A bearing bumper 58A mayalso be provided to reduce bending, as is described further below.

In use, bearing loads exerted on inner hub 76 are transferred to outerring 68 via struts 40, as follows. In general, bearing loads generatedby engine thrust and transient dynamic events, such as blade-off eventsor bird strikes, are experienced mainly at bearing set 57 (bearing 58typically contributes little additional loading in such events) whichare passed into the inner hub 76 at its leading edge. The inner hub,with its semi-monocoque design, reacts to the applied loads internallyas tension/compression and shear, as described above. The bearing loadis passed mainly through the leading edge 40A of the strut 40 incompression or tension to the mount pads 74. For reasons describedbelow, the mount pads 74 are located at (or near) the centroidal axis CAof the strut 40 cross-section.

In use, engine inertia loads are also exerted on the splitter 42 by theremainder of the engine connected thereto via the gas generator case,and these are transferred to outer ring 68 via struts 40. In general,engine inertia loads enter the intermediate case 46 via the splitter (towhich the gas generator case is attached) and are reacted in the rearouter portion 40D of the strut 40 as a compression or tensile load.These loads tend to bend the strut and torque box and thus are reactedinto the structure of strut 40 by the reaction of torque box 41converting the load into a shear which stiffener 94 transmits as atension or compression into the rear of the strut. The torque boxes 41will now be described in more detail.

The torque boxes 41 are hollow closed cells formed between the struts40, splitter 42, and stiffener 94. As will become apparent below, torqueboxes 41 are somewhat similar in purpose and function to the torque boxpresent in an aircraft wing, although here the construction is analogousto an aircraft wing wrapped into a cylinder. The rear stiffener web 94,it will be seen, is analogous to the spar of this cylindrical wing. Thetorque boxes 41 “convert” loads applied to one or more struts (forexample, a bending moment and a transverse shear) into a balanced shearflow in the cell, which may then be “communicated” to and reacted byadjacent struts, as will now be described.

Referring to FIGS. 15, 16 a and 16 b, a load, such as a bending moment,in one direction on one strut 40 will be communicated by the torqueboxes 41′ to the two adjacent struts 40′, which will in turn of coursereact the force, thus tending reduce the effect of the applied load onthe first strut by transferring a reactionary component to the adjacentstruts. In this manner load sharing is achieved. (Though only theinteraction of three struts is shown in FIG. 16 b for descriptionpurposes, it will be understood that struts 40′ likewise communicateexternal and internal loads to their adjacent neighbors via theirrespective torque boxes, and thus external and internal loads are thusredistributed around the structure among the struts 40.) Referring stillto FIG. 16 a, and as will be discussed in more detail below, a torsionalload applied to torque box 41 (represented by the circular stippledarrow), such as that applied by the weight/inertia of the gas generatorattached to the splitter, is also reacted by the torque box 41, in thiscase preferably mostly as a shear force, which is passed to strut 40 asan in-plane load at least partially by a shear (represented by thestraight stippled arrow) passed through the shear transmitting joint 94Afrom web 94 to strut 40. The stiffener ring 43 helps to distribute theinertia loads more uniformly to the torque boxes 41. The torque boxarrangement and structure therefore both helps distribute loads amongadjacent struts as well as convert torsional and bending loads intoshear, which can then be transmitted as substantially pure (preferably)compression or tension in struts 40.

Therefore, since the struts are inherently connected, any tendency fordisplacement of one strut is inherently reacted and balanced through thetorque boxes by adjacent struts, which not only redistributes the loadbut also substantially reduces the amount of bending forces on thestruts, even during transient dynamic events such as bird strikes. Thissignificant reduction of bending forces which permits the use ofthin-walled structures of the struts of the present invention, since theabsence of plate bending permits substantial reduction incross-sectional thickness in the casing and struts relative to the priorart.

Referring still to FIG. 16 a, the in-plane loads transferred from torquebox 41 to strut 40 will thus load the aft portion 40D of the strut 40 intension or compression (depending on load direction) and this internaltensile or compressive load is then carried by the aft portion 40D ofthe strut 40 to the outer ring 68 and ultimately the engine mount 74.The shape of the strut 40 is used to divide the bearing loads from theinertia loads. In particular, the bend or kink 40E in the aft portion40B of the strut 40 reduces the axial stiffness of the strut 40 whichthus creates two separate load paths for the loads generated in theengine (i.e. one for bearing loads and one for inertia loads, asdescribed above). The kinked shape of the strut 40 interrupts the loadpath to the inner hub, which thereby impedes the transfer of loads fromthe splitter to the hub. This simplifies load transfer as will asbeneficially reducing bending on the strut, which thereby permits athin-walled strut structure to be employed. Referring to FIGS. 18 a and18 b, since prior art struts were required to react bending forcestransmitted thereto, the prior art struts required thick enough sections(FIG. 18 b) to provide the appropriate bending strength. In the presentinvention, however, the reduction, or more preferably negation, ofbending of strut 40 permits the use of sheet metal struts (FIG. 18 a)which are of course much lighter than the prior art.

As described above, the engine mounts are preferably positioned along(or as close as is possible) the centroidal axis, thereby negating (orreducing to a manageable level) the bending moment applied tointermediate case 46 as a result of the tensile/compressive loads passedto the intermediate case 46 from struts 40. In this manner, bending isreduced on intermediate case 46 and struts 40, further enhancing theopportunity to make full advantage of the semi-monocoque and thin-walleddesign of the case and struts to thereby maximize structural efficiencyand minimize weight. The structural efficiency of the semi-monocoquestructure of the inner hub 76 and outer ring 68 is thereby improved andenhance by the use of the struts 40 of the present invention, andalthough these components may be employed individually with advantage,the use of two or more, and preferably all three together provides yetfurther advantages and benefit by the intrinsic co-operationtherebetween which may be obtained.

It should be noted that, as described above, the balanced shear flow,induced in the torque boxes 41 as a result of a torsional load, isreacted by the struts 40 predominantly as shear load at thesplitter/strut joints (42A, 94A). Thus, there is a substantial absenceof tensile loads at these joints, which advantageously permits the useof fillet welds to provide joints 42A, 94A. Also, due to the relativelylong length of these joints, and loading sharing among the plurality ofjoints in the overall structure (i.e. on the plurality of torque boxes),the shear stresses on the joints are relatively low, thus furtherallowing a reduction of the thickness the strut and torque boxcross-section. Very thin gauges of sheet metal may thus be used.

Advantageously, the struts may be designed to act as a load “fuse”limiting the allowable load transmitted to the mount by theircompressive capability. (It will be understood that when a sufficientcompressive load is applied to the thin-walled strut, the strut willcollapse). For example, the strut may be designed to collapse when acertain threshold load is experienced (e.g. a significant big strike) tothereby limit the amount of load (and therefore damage) which istransferred to the aircraft in such an event. In this example taken inthe context of the preferred embodiment above, when the thresholdbearing load is applied by the inner hub to the strut, the leading edgeis designed (i.e. by virtue of its thickness, etc.) to collapse undersuch event loads, thereby absorbing energy by plastic deformation ratherthan transferring it to the engine mounts and aircraft. In design, themaximum allowable load to be transferred by the strut would bedetermined, and then a strut configuration is determined that wouldcollapse or otherwise structurally fail upon the application of thismaximum load, or a larger load, and thereby limit the load transfer tothe engine mounts.

Referring again to FIG. 13, the bearing bumper 58A can be provided toassist in improving the stiffness of inner hub 76. For example, sizableasymmetric bearing loads are applied to inner hub 76 during medium-sizedbird strike events, for example, which tend to cause bending in theengine shafts, which tend to distort the bearing housing, and thusbearing seat 58. The bumper 58A is a leg or stop-type device which isprovided with a small clearance (not shown, as the scale of FIG. 13. istoo small to indicate this feature) between the bumper 58A and thebearing seat 58 (or bearing or other appropriate surface). The clearancepreferably corresponds to the amount of allowable deflection desired insuch an event (e.g. 0.005″, for example). If a larger deflection isforced, the bumper will assist the bearing seat 58 (or whatever surfaceis opposed by bumper 58A) to resist such deflection. This simple devicetherefore permits the rear portion of the inner hub 76 (i.e. the portionsupporting the bearing seat 58) to be substantially thinner, since theinner hub 76 thickness does not need to react these bending forces anddeflections alone. This therefore also helps unload the bottom and rearportion of the strut 40, so that the inner hub 76 and bearing seat 58can be thinner, and weigh less.

Although the individual weight savings achieved by each aspect of thepresent invention may be insubstantial when considering larger turbofanengines, in the case of very small turbofan engines (e.g. 2000 poundsthrust and under), these accumulations of small weight savings result ina significant weight savings.

The invention provides a multi-faceted structure which seeks to forceout-of-plane loads (e.g. bending loads) back into plane, and balancestensile and compressive loads with shear panels to thereby create equaland opposite shear flows in adjacent panels.

In this application, “thin wall” means sheet metal type thickness,wherein “thin” is interpreted relative to the applied loads, such thatthe thin wall is substantially incapable of reacting applied bendingforces in plate bending.

While the above description addresses the preferred embodiments, it willbe appreciated that the present invention is susceptible to modificationand change without departing from the scope of the accompanying claims.For example, while described in respect of an application to very smallturbofan engines, some benefits may be attained in larger turbofan orother gas turbine engines in applying the principles of the presentinvention. Though the use of certain materials and manufacturing methodshave been disclosed as preferred, other materials and methods may besubstituted without departing from the present invention. The cases neednot be integrated as described to achieve benefits of the presentinvention. Likewise the struts need not necessarily be hollow in allembodiments, nor need they comprise a single “cell” as described above,but may have multiple cells defined therein (see FIG. 19). As shown inFIG. 16 c, the torque box may comprise more cells. The torque box neednot be comprised of the splitter itself, but may be an additionalstructure which may be inside the splitter, or elsewhere. Although asingle strut is preferred for transfer of both bearing and inertialoads, multiple struts (e.g. an upstream and downstream strut pair) maybe sued). The semi-monocoque shear panels in ring 68 and hub 76 need notbe rectangular or regularly sized. Still other modifications will beapparent to those skilled in the art which will fall within the scope ofthe invention intended by the inventors, and the appended claimstherefore are not intended to exclude such modifications.

1. A casing for an aircraft turbofan bypass engine, the casingcomprising: a thin-walled sheet metal annular case adapted to encirclethe engine and having a plurality of engine mounts thereon adapted tomount the engine to an aircraft, airflow through the engine in usedefining upstream and downstream directions; a plurality of thin-walledsheet metal struts connected to an annular leading edge tip of asplitter separating a core flow passage from a bypass flow passage, thestruts having a leading edge extending uninterruptedly between an enginebearing support structure and the case, the strut leading edgesconnected to the engine bearing support structure axially upstream ofwhere the strut leading edges connect to the case thus angling the strutleading edges rearwardly as they extends radially outwardly, the enginebearing support structure disposed inwardly of the core flow passage,the strut leading edges disposed upstream of the splitter leading edgetip such that the strut leading edges provide an uninterrupted primaryload path from the engine bearing support structure to the case fortransfer of thrust loading from engine bearing support structure to theengine mounts on the case, the struts each having a centroidal axisdefined along a locus of centroid positions for a plurality of strutsections along a length of the strut, wherein the engine mounts areaxially and circumferentially positioned on the case to substantiallycorrespond with one of said strut centroidal axes to thereby minimizebending loads in the case as a result of thrust loads transferred by thestruts to the engine mounts.
 2. The casing of claim 1 wherein the caseis an intermediate case portion of the casing.